1. Field of the Invention
The present invention relates to the cooling of components exposed to hot gas atmosphere and, more particularly, pertains to internally cooled gas turbine engine airfoil structures.
2. Description of the Prior Art
Referring to FIG. 6A, conventional internally-cooled turbine rotors typically comprise a disc 2 supporting a plurality of circumferentially-spaced turbine blades 1 having at least one internal cooling channel 5 defined therein, the cooling channel having an entrance opening 6. Often, there is more than one such channel, and FIGS. 6B and 6C shows three such channels, for example, labelled X, Y and Z for convenience. The root 3 of each of the blades is positioned in a slot in the disc. Defined between the blade and the disc is a cooling air channel or pocket 4 which communicates with the blade internal cooling channel via the entrance 6. In use, the cooling air pocket is fed with cooling air, for example from a tangential onboard injector (TOBI) or other means, and from there flows through the entrances 6 and into the internal cooling channels 5 for the purpose of cooling the blade.
The high rotational velocity of the turbine rotor relatively to the cooling air supply makes it generally difficult to feed the blade internal cooling passages. Air must be redirected several times, at several angles which are almost normal to each other, which is exceedingly difficult to do efficiently in high speed rotating machinery. Although the TOBI provides a partial solution, as depicted in FIG. 6B the air entering the cooling air pocket tends to generate a considerable re-circulation vortex inside the pocket, the vortex being caused by air entering the pocket at an angle (due to relative rotation of the disc) and then impacting and being redirected by a “downstream” first side of the pocket (the downstream side of the pocket is depicted along the bottom of FIG. 6B) and thus guided therealong to the back of the pocket. The difficulty in directing air results in an uneven cooling flow split among the various blade entry cooling passages. Referring to FIGS. 6B and 6C, the uneven cooling flow split tends to result in a larger percentage of the overall cooling flow entering passage Z (represented generally in FIGS. 6B and 6C by the disproportionate arrow sizes), which thereby reduces the efficiency of the cooling achieved through passages X and Y.
EP 1251243, published on Oct. 23, 2002, speculates that an air distribution problem between passages is caused by a low pressure region in the centre of the re-circulation vortex (which pressure is generally lowest at the point corresponding to the location of passage Y), and thus teaches installing a fence on the under-surface of the blade root to extend into the pocket and disrupt the swirl of cooling air. The U-shaped metal sheet EP 1251243 appears to act as a flow splitter, which attempts to break the vortex structure of the coolant flow, to thereby prevent the formation of low pressure zone inside the cooling air channel.
Though EP 1251243 may offer some improvement, there is still a need for an improved means for supplying a coolant air flow to internally cooled airfoil blade which will provide a better pressure and flow distribution between cooling passages with the blade.